The invention relates to internally cooled aerofoils.
In particular it concerns the design of the internal structure of an internally cooled turbine aerofoil for a gas turbine propulsion engine.
The surface temperature of the aerofoils in the high pressure turbine of a gas turbine engine can reach the melting point of the alloy of which they are constructed. Hence heat is extracted by passing cooling air through passages formed in the interior or in the walls of the aerofoil. Known methods of casting internally cooled turbine aerofoils use ceramic cores that have been manufactured in a die. Consequently the shapes of the internal wells, or webs, that define the internal cooling passages are constrained by the requirements for die draw/withdrawal angles. The webs of the internal structure of aerofoils containing multiple cooling passages have been made thinner either in the middle or on one side of the blade, depending on the complexity of the die. Examples of internal structures of this kind are illustrated at FIGS. 1 and 2.
These designs suffer a drawback in that the shape of the internal webs forming the internal support structure for the aerofoil outer wall have been optimised for the ease of manufacture of the ceramic core or the aerofoil rather than for maximum aerofoil life.
Japanese Patent No. JP 60135605 entitled “Turbine Blade Cooling” published in 1985 disclosed an internally cooled turbine blade in which internal walls, webs, partitioning flow passages in the interior of the blade are made into convex forms. This convex form reduces the cross-sectional area of the flow passage, compared to the same blade in which the partition walls are not convex, resulting in increased flow speed in the passage and increased coefficient of heat transfer by convection. The chordal extent of the convex webs appears substantially constant along the length of the flow passage. Notwithstanding increased heat extraction, increased mass in a rotating component at high radius will adversely affect the creep resistance of the blade and this may be a limiting factor in blade life. This might be the case where engine weight is important as in an aircraft propulsion engine.
Japanese Patent No. JP 61205301 entitled “Gas Turbine Blade” teaches an arrangement to dissipate the unevenness in the temperature of a gas turbine blade in which a fluid cooling passage is formed. The sectional area of the cooling passage is smaller in the mid-height part of the blade than in the root and tip regions. The partition walls between flow passages are provided with projecting parts, or ribs, to enhance heat transfer. This arrangement is designed only with regard to the distribution of temperature in the external turbine gas flow.